In recent years, aircraft manufacturers have developed aircraft designs and aircraft fabrication methods that make greater use of carbon fiber composite materials and the like (“composite materials” or “CFCM”), such as graphite/epoxy and carbon fiber reinforced plastic (“CFRP”). Composite materials are significantly lighter than traditional aircraft materials (e.g. aluminum, titanium, steel and alloys of these), and can provide high strength with low weight, allowing lighter, more fuel efficient aircraft. In some newer aircraft, for example, the majority of the primary structure, including the fuselage and wing, is made of composite materials. By volume, some new aircraft can be about 80% composite materials.
Since composite materials have different characteristics than some traditional aircraft materials, new facilities, equipment and handling methods have been developed. This includes fabrication methods for wing structures. A typical wing structure of an aircraft includes one or more main spars, which extend from the root of the wing to the wing tip. These wing spars typically taper from the root to the tip of the wing, and have a plurality of ribs attached along their length. These ribs include interior ribs, located between the main spars, and leading edge and trailing edge ribs. The ribs are oriented generally perpendicularly to the wing spars, and together define an outline that corresponds to the cross-sectional shape of the wing at each rib location.
Attached to the wing ribs are skin panels that provide the finished shape and aerodynamic contour of the wing. The portion of the wing structure from the front main spar to the rear main spar is referred to as the wing box. In composite aircraft, wing skin panels that cover the wing box are typically much thicker than the skin panels for the leading or trailing edges. However, the wing edge skin panels attach to the wing box panels. In the case of aircraft with moveable slats on the leading edges of the wings, a fixed leading edge panel is attached to the wing along the leading edge beneath the slat.
In order to provide a suitable connection and joint for fixed edge skin panels in a composite aircraft wing, a variety of connection methods have been developed. Some existing methods use devices such as splice straps, joggled panels, wiggle plates or secondarily bonded fillers to facilitate this joint. These can be heavy, complex and expensive to make. Some approaches to this issue add extra processing steps in order to ensure the integrity of the joint.
The present disclosure is directed toward one or more of the above issues.